1. Field of the Invention
The present invention relates generally to gas turbine engine, and more specifically to a large highly tapered and twisted and thin turbine rotor blade with multiple impingement near wall cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, such as a large frame heavy-duty industrial gas turbine (IGT) engine, a hot gas stream generated in a combustor is passed through a turbine to produce mechanical work. The turbine includes one or more rows or stages of stator vanes and rotor blades that react with the hot gas stream in a progressively decreasing temperature. The efficiency of the turbine—and therefore the engine—can be increased by passing a higher temperature gas stream into the turbine. However, the turbine inlet temperature is limited to the material properties of the turbine, especially the first stage vanes and blades, and an amount of cooling capability for these first stage airfoils.
The first stage rotor blade and stator vanes are exposed to the highest gas stream temperatures, with the temperature gradually decreasing as the gas stream passes through the turbine stages. The first and second stage airfoils (blades and vanes) must be cooled by passing cooling air through internal cooling passages and discharging the cooling air through film cooling holes to provide a blanket layer of cooling air to protect the hot metal surface from the hot gas stream.
As the turbine inlet temperature increases with higher efficiency engines, later stages of the turbine rotor blades will require cooling. The latter stages of blades are also large blades with high amounts of taper and twist. The fourth stage turbine rotor blade can be over three feet in spanwise length and is too thin for most types of internal cooling circuits. For a large turbine rotor blade, cooling holes are drilled radial holes from the blade tip to the root section. Limitations of drilling a long radial hole from both ends of the airfoil increases for a large and highly twisted blade. A reduction of the available airfoil cross sectional area for drilling radial holes is a function of the blade twist. Higher airfoil twist yields a lower available cross sectional area for drilling radial cooling holes because a straight path from the tip to the root is not available. Cooling of the large and highly twisted blade by this manufacturing process will not achieve the optimum blade cooling effectiveness. U.S. Pat. No. 6,910,864 issued to Tomberg on Jun. 28, 2005 and entitled TURBINE BUCKET AIRFOIL COOLING HOLE LOCATION, STYLE AND CONFIGURATION shows a profile view of a prior art large rotor blade cooling design with drilled radial cooling holes as described above.
Alternative designs to the radial cooling channels for these large and highly twisted turbine rotor blades have been proposed such as the use of multiple pass serpentine flow or multiple radial channels with pin fins for cooling. However, producing a ceramic core to achieve an acceptable casting yield for a large tapered and twisted blade has not been found. Ceramic cores must be made into more than one piece which leads to core shifting during the casting process or from core pieces breaking such that the cooling circuit is not completely formed.